Trends in Flight-Operated Small-Satellite Propulsion Technologies

Abstract

The development and execution of prospective inner and outer space missions require focusing on the use of many small space vehicles operating in swarms with multiple informational, navigational, and mission-oriented interactions among themselves. Such missions involve providing communication and surveillance services, facilitating distributed material production in space, and conducting research expeditions to explore the resources and environments of new worlds. The cornerstone technology for operating distributed space systems is propulsion. Among a range of propulsion technologies—from using pressurized cold gases to implementing laser beams to generate thrust—certain methods stand out for application in small spacecraft. This paper provides a summary of space-operated propulsion, emphasizing the reasons for the more frequent adoption of one technology over another. The discussion on propulsion trends is complemented by examining the physical, engineering, production, operational, and societal rationale behind these choices. The findings reinforce the trend toward transitioning to fully electric satellites. This review serves as a means for reevaluating global propulsion trends and guiding the future development of inner and outer space propulsion-assisted economies effectively.

1. Introduction

Recent decades have witnessed profound transformations in the aerospace sector driven largely by the proliferation of small satellites [1,2,3,4]. These compact space vehicles, which typically weigh less than 500 kg, have revolutionized various fields of space exploration, including communication, remote sensing, and scientific research. Their affordability [5], adaptability [6], the resilience of the systems based on them [7,8], and their swift deployment capabilities [9,10] have positioned them as essential tools for governments, corporations, and research entities [11].

The evolution of small satellites has intensified scrutiny on their propulsion systems, especially in the last decade [12,13,14]. Propulsion plays a pivotal role in ensuring that the satellites achieve and maintain their designated orbits, execute maneuvers, and fulfill mission objectives efficiently [15,16,17]. Selecting the right propulsion technology for small satellites is a multifaceted task that demands thorough consideration of factors such as mission duration, orbital requirements, and maneuvering demands [18,19].

Propulsion systems are especially important for proliferated space systems, which are seen as the key means for the development and prosperity of space economies [20,21,22]. These systems should enable not only precise orbit maintenance but also allow for complex maneuvers necessary for different applications [23]. Additionally, these propulsion systems should be capable of being serviced in-orbit [24,25,26,27,28]. They should also control the thrust vector direction, enabling precise adjustment of direction and magnitude of thrust [29]. This feature enhances the overall performance and reliability of small satellites, ensuring they can effectively carry out their intended functions [30].

However, the use of propulsion systems on small satellites is associated with a number of fundamental limitations due to their small dimensions. The most significant limitation is the available electrical power. For example, for a CubeSat-class satellite in the 3U format located in low Earth orbit and equipped with a deployable solar power system consisting of seven solar panels with an energy conversion efficiency of about 30%, the peak-generated power per orbit is approximately 59 W [31]. Given a mass of such a satellite of around 4 kg, this allows estimating the upper limit of specific power for the energy system at about 15 W per kilogram. Note that there are experimental NASA systems capable of providing up to 200 W of power to a 3U satellite [32]. However, such solutions require complex, large-scale deployable structures, which have not been widely adopted due to constraints on mass, cost, and reliability. In contrast, full-sized spacecraft have significantly higher energy capabilities, achieving specific powers of about 25 W per kg, while the total generated power ranges from 10 to 20 kW [33]. This disparity in energy supply underscores the need for specialized solutions when selecting a propulsion system for small platforms.

Another challenge faced by small satellites involves difficulties in organizing the thermal stabilization of all structural elements and integrating the propulsion system with the satellite itself. Limited dimensions make it challenging to accommodate traditional heat dissipation systems used in larger satellites. Despite these challenges, successful engineering solutions have been implemented within several missions, exemplified by the Lunar IceCube mission, where a system of several small radiators was developed for component thermostabilization [34]. The integration of engines and satellite systems primarily hinges on minimizing losses in actuation systems. Small satellites demand compact and energy-efficient actuation systems. Similarly rigorous standards apply to pressure, temperature, flow-rate sensors, etc., ensuring appropriate control over the propulsion system’s operations. Contemporary solutions in this domain include specialized microsystems alongside algorithms for state estimation based on indirect parameters [35,36,37].

Presently, a broad spectrum of propulsion technologies exists. Nonetheless, a comprehensive analysis of their applicability aboard small satellites remains insufficient. This highlights the critical importance of reviewing the collective experience gained by various research entities concerning the implementation of such systems on small spacecraft.

This paper systematically reviews the trends and prospects of propulsion technologies for small satellites. Starting with an examination of the classification of space missions and highlighting the importance of propulsion for modern small-satellite missions, the discussion proceeds to analyze flight-operated propulsion systems, their merits and limitations, and their implications for small-satellite operations. The propulsion systems included in this review were chosen based on two primary criteria: the propulsion systems are specifically designed for operation aboard small satellites, and these systems have already been tested in space. The selected propulsion systems are categorized, according to the classical classifications, into three groups: chemical, electric, and alternative. Space tethers were excluded from this review because they do not appear suitable for dynamic maneuvering aboard small satellites. Experimental evidence and theoretical insights are emphasized, focusing on critical factors related to five rationales—physics, engineering, production, operational, and societal—to provide a structured evaluation framework revealing potential directions for prospective propulsion technologies applicable to small satellites for future research and development efforts.

This review serves as a tool for reevaluating global propulsion trends and guiding the future development of inner and outer space propulsion systems to assist in the effective development of space-related economies. Section 2 introduces the classification of space missions focused on small satellites, stressing that propulsion technologies constitute the most critical component of these missions. Section 3 discusses different propulsion technologies that have been deployed on board small satellites in space. Finally, Section 4 explores the various types of rationales driving the development, study, and operation of small-satellite propulsion systems, along with current trends and future perspectives in the field.

2. Space Missions of Small Satellites

Modern space missions based on small satellites demonstrate considerable diversity in terms of purpose, orbital characteristics, and operational requirements. This highlights the necessity of introducing a classification framework for missions conducted by such platforms. It is proposed to categorize them into subclasses according to the following criteria: orbital destination, formation architecture, mission duration, functional objectives, and propulsion-related attributes. Figure 1 visually represents an expanded classification of space missions using small satellites.

The orbital destination defines basic energetic and temporal characteristics of the mission. Low Earth orbits (LEOs) are characterized by substantial resistance from residual atmosphere. Consequently, a propulsion system is required onboard to maintain altitude. The demand for characteristic velocity change (ΔV) increases exponentially with decreasing altitude to compensate for atmospheric drag. At altitudes between 350 and 450 km, annual ΔV amounts range from tens of meters per second, annually. Medium Earth orbits (MEOs) lack any remaining atmosphere but still require periodic adjustments to preserve navigation-related parameters. Thus, the necessary ΔV here depends heavily on the complexity of planned maneuvers in each mission. Geosynchronous orbit (GEO) demands extremely accurate positioning (within ±0.05°) and, consequently, requires significant ΔV (50 m/s annually) to counteract drifts from assigned positions. Highly elliptical orbits with large eccentricity necessitate regular corrections to maintain consistent periods of revolution. Interplanetary missions and deep-space operations impose the strictest requirements regarding ΔV increments, typically requiring more than 1 km/s.

The formation architecture dictates the scale of the system—from single satellites to megaconstellations comprising thousands of units. Single satellites critically depend on autonomous orbit management. Formations involving 2–10 satellites require precision in relative positional maneuvering among themselves. Constellations and megaconstellations prioritize standardized systems with highly reliable propulsion units.

Mission duration influences the choice of propulsion system type. Short-term missions (<3 months) may rely on passive means or low-specific-impulse propulsion systems (chemical rockets or cold gas thrusters). Mid-term missions (12–24 months) impose stricter specific-impulse requirements because of limited propellant storage volume in small satellites, making EP preferable. Long-term missions (over three years) demand even greater specific impulses and durability from propulsion systems. Alternatively, propelantless methods like solar sails might also be employed.

Functional objectives determine unique requirements for maneuver execution. Scientific missions, such as studying cosmic radiation effects on organisms in interplanetary space, generally involve standalone satellites requiring precise positioning and minimal disturbance for experiment integrity. Conversely, commercial missions often deploy groupings of satellites, for example, the HawkEye 360 mission to identify geolocation sources of radiofrequency emissions, emphasizing economic feasibility and reliability of propulsion units.

In addition to external factors, another subclassification is tied to internal factors, specifically related to specific impulse (I_sp) and thrust performance of propulsion systems. Specific impulse governs the effectiveness of propellant usage. As highlighted earlier, for missions with restricted mass and volume of propellant storage and supply systems, especially critical in CubeSats, achieving high specific impulse is crucial to meeting cumulative ΔV needs. Ion thrusters like BIT-3 and NPT30-I2 exhibit very high specific impulses (2000–3000 s), which are suitable for missions requiring extensive ΔV, including interplanetary travel, orbit transitions, and long-term orbit maintenance. However, their further miniaturization for <6U platforms remains challenging due to requirements for cathode-neutralizers and complex ion-optical systems. Flights of Lunar IceCube (2022) and LunaH-Map (2022) [12] validated the functionality of ion thrusters onboard 6U CubeSats but revealed certain problems addressed later in subsequent chapters. Electrospray thrusters similarly feature high specific impulses (1000–5000 s), e.g., TILE and S-iEPS models, although their application is limited by contamination risks from droplets of propellant and emitter degradation over extended operation that can be critical for resilient space systems.

Thrust, another key parameter, determines the time needed to execute maneuvers and the ability to overcome disturbing forces. To promptly adjust orbits in LEO, it is desirable to deliver a thrust of at least 1–5 mN to offset atmospheric drag. Resistojets, exemplified by AQUARIUS and ARM-A, provide 1–20 mN of thrust, rendering them effective for quick orbit-correction and rendezvous tasks. Their simplicity and robustness were proven through successful flights, notably EQUULEUS (2022), which pioneered water-based resistojet-assisted orbit correction beyond LEO. Electrodeless plasma thrusters, such as BDEPT and Maxwell, could also serve similar purposes. BDEPT flight operations on HORS 1 (2023) verified its 2U model’s capability to generate up to 10 mN of thrust. On the other hand, previously mentioned high-specific-impulse engines (ion and electrospray types) suffer from much lower thrust levels (0.1–100 μN), limiting their suitability for rapid corrections. Nonetheless, their ultra-low thrust suits precision-positioning tasks well. Pulsed plasma thrusters, despite offering comparable thrust levels, benefit from simpler feeding mechanisms but fall behind in specific impulse.

Thus, analyzing various small-satellite mission requirements alongside existing propulsion characteristics leads to several conclusions. The broad spectrum of tasks performed by small satellites, ranging from massive communication constellations to unique scientific missions, necessitates diverse types of propulsion systems to be employed. Thruster selection should consider multiple factors simultaneously: specific impulse, thrust, accuracy of control, mass/volume constraints, total impulse, and operational lifetime. However, theoretical evaluations alone cannot substitute validation via actual space operation, where space effects arise that are overlooked in terrestrial testing. Therefore, systematic analysis of flight-experience-derived insights forms the basis for informed decision-making regarding optimal thruster choices—a subject further explored in the Section 3 of this study.

3. Space-Operated Small-Satellite Propulsion Technologies

3.1. Chemical Propulsion

There are three main types of chemical propulsion—the monopropellant rocket (see Figure 2), the bi-propellant rocket (see Figure 3), and the hybrid rocket (see Figure 4) engines. The operation of chemical propulsion depends on the type of propellant being used. For single-component propellant engines (monopropellant rocket engines), thrust is generated through thermal decomposition and acceleration of the products in the nozzle. In this case, the propellant comes into contact with a catalyst inside the combustion chamber, leading to chemical breakdown, heat release, and formation of high-temperature gases. Under high pressure, these gases exhaust through the nozzle, generating thrust [38,39]. In two-component systems (bi-propellant rocket engines), it is required to use fuel and an oxidizer, which are stored onboard either in liquid or gaseous forms. These components mix in a specified ratio and ignite spontaneously when brought together, or they require external ignition devices. Combustion produces hot gas, which, under high pressure, is exhausted through the nozzle, generating thrust [38,39]. Chemical engines also encompass systems that utilize propellants in different states (hybrid rocket engines) [40].

Table 1. Flight-operated chemical rocket engines.

PS	Entity	Propellant	P, W	T, N	Isp, s	It, kN·s	Size	Mass, kg	Missions (Year)	NORAD ID	Ref.
MPS Ø	UTokyo, Tokyo, Japan	H2O2	-	0.5	<80	-	-	-	Hodoyoshi-1/Hodoyoshi-3 (2014)	40299/40015	[49]
EPSS C1K ●	NanoAvionics, Vilmius, Lithuania	ADN	7.5	<0.3	214	0.4	1.3U	1 *	Lituanica-2 (2017)	42768	[50]
Steam Propulsion ▲	The Aerospace Corp., Chantilly, VA, USA	H2O	12	0.004	70	-	<1U	-	Aerocube 7 (2017), Aerocube 10 (2019)	40966, 44485	[42]
PM200 ●	Dawn Aerospace, Delft, The Netherlands	N2O + C3H6	12	0.5	285	0.85	1U	1.1 * 1.4 **	Hiber-4 (2021)	47541	[51]
HYDROS-C ▲	Tethers Unlimited, Bothell, WA, USA	H2O	<25	>1.2	<241	<3.38	2U	2.7 **	PTD-1 (2021)	47482	[44,52]
LFPS ▲	NASA MSFC, Huntsville, AL, USA	ASCENT	<47	0.1	<200	<3.5	2.4U	5.5 **	Lunar Flashlight (2022)	54697	[45,52]
Monopropellant CubeSat System ●	Stellar Exploration, San Luis Obispo, CA, USA	Hydrazine	-	0.25	200	-	-	-	NASA Capstone (2022)	52914	[53]
ArgoMoon Hybrid MiPS X	ECAPS, Solna, Sweden	LMP-103S	20	100	190	0.783	1.3U	1.43 * 2.07 **	ArgoMoon (2022)	55907	[54]
HyPer ▲	The Aerospace Corp., Chantilly, VA, USA	H2O2	-	-	<124	-	0.25U	-	Slingshot-1 (2022)	52947	[55]
HAN-based propulsion unit Ø	Hunan Hangsheng Satellite Technology, Changsha, China	HAN	-	-	-	-	-	-	Jinta (2023)	56169	[55]
CubeDrive 0.8U ●	Dawn Aerospace, Delft, The Netherlands	N2O + C3H6	15	0.49…1.35	<248	0.4	0.8U	1.051 * 1.250 **	SC1 (2024)	62388	[56]

* Dry mass. ** Wet mass. ^▲ PS distinct operating parameters, which are available in open information sources. ^● PS demonstrated in orbital flights, but data on its characteristics varies in different sources. ^Ø PS demonstrated in orbital flights, but there is no open data on its characteristics. ^X Propulsion systems that have been launched into space but have failed to successfully demonstrate their functionality in orbital conditions for various reasons.

presents flight-operated chemical propulsion units suitable for small satellites. All mentioned engines typically have analogous versions within the same series, featuring larger sizes, higher thrust, and greater power outputs [36,37,38,39,40,41,42,43,44,45,46,47,48,49,50,51,52,53,54,55,56].

Analysis of missions using chemical propulsion (see Table 1) reveals the challenges associated with its widespread adoption on platforms of a size less than 6U [36,37,38,39,40,41,42,43,44,45,46,47,48,49,50,51,52,53,54,55,56]. The main reason lies in the low propellant storage density, constrained by the nature of propellants and the dimensions of satellite subsystems, combined with a low specific impulse (Isp), leading to restrictions on its applicability for short-term missions [39,40,57].

Bi-propellant propulsion systems rarely achieve a specific impulse exceeding 285 s [51], whereas modern electric propulsion systems reach values above 1500 s. Given equal masses of propellant for a chemical rocket and for electric propulsion, the latter would provide a characteristic velocity increase of more than five times. For CubeSats, where the propellant mass usually accounts for no more than 20% of the total vehicle mass, this disparity makes chemical propulsion units impractical for missions lasting longer than six months [18,19]. Exceptions include tasks requiring instant high thrust (>100 mN), such as rapid orbital correction after separation from the launcher or fast maneuvers to avoid space debris [30,31]. Additionally, propellants used in such systems often turn out to be toxic, complicating ground testing and necessitating careful selection of materials for propellant tanks [45,46,53].

Green monopropellant systems based on hydrogen peroxide (H₂O₂) and ammonium dinitramide (ADN) occupy a special niche [42,50]. The HyPer installation (Aerospace Corporation, Chantilly, VA, USA), which underwent 34 successful tests aboard the 12U Slingshot-1 satellite (2022), confirmed operational reliability with reduced toxicity compared to hydrazine [47,48]. Nevertheless, energy limitations persist: with I_sp = 124 s and a propellant mass of 0.5 kg on a 19 kg platform, the achievable characteristic velocity reaches only 21.5 mm/s, which is insufficient for correcting orbits in LEO without continuous atmospheric drag compensation [47,48,51].

Hybrid architectures combining chemical and electric propulsion represent a promising path to overcome practical limitations of chemical rockets [39,40]. The most successful implementation has been the hybrid HIPS system on the CAPSTONE mission (2022), where eight monopropellant hydrazine rockets provided trajectory corrections toward the Moon and precise positioning on halo-orbit. Such a task would be unfeasible for purely electric systems given a solar array output below 100 W, thus defining a prospective niche for future development of chemical systems.

It is important to pay particular attention to failed missions, which highlight systemic scaling issues. The failure of the LFPS engine aboard Lunar Flashlight (2022) was caused by solid particles blocking the fuel line [46]. This problem becomes exacerbated during the miniaturization of fuel feed channels under strict mass limits for filtering components. Similarly, two out of four satellites in the Hiber constellation (2021) with bi-propellant PM200 engines experienced failures linked to uneven nitrous oxide vaporization in microgravity conditions, resulting in stoichiometric mixture imbalance. These cases emphasize how chemical processes stable at full-sized spacecraft scales become sensitive to technological peculiarities when transitioning to smaller vehicles [39,40,57].

In conclusion, bi-propellant and monopropellant chemical engines remain competitive for addressing short-duration LEO missions requiring initial orbital adjustment post-separation and tasks demanding high-thrust levels (>50 mN), such as towing space debris or fast maneuvering. Hybrid architectures stand out as they enable broader mission scopes, since chemical thrust complements electric propulsion systems for performing impulsive maneuvers while maintaining overall energy efficiency.

3.2. Cold Gas

Cold gas thrusters are among the simplest and most reliable in operation, and they are widely used in space systems for orientation and stabilization, orbit correction, and other purposes. The operational principle of these thrusters is based on utilizing compressed gas energy. The schematic diagram of the cold gas thruster is depicted in Figure 5. In its simplest configuration, high-pressure gas is stored in a tank. Upon command, the valve connecting the fuel tank and nozzle opens, allowing the gas to pass through a filter designed to eliminate foreign particles before entering the nozzle. The gas exits through the nozzle, generating reactive thrust [57].

More complex systems include two tanks: one with liquid propellant and another intermediate tank that maintains part of the propellant in gaseous form. The propellant exhausts through the nozzle, with a hydraulic connection line to the tank, in which the propellant is in the gaseous state. Then, this tank is refilled with the propellant stored in a tank in which it is in liquid form [57]. Typically, the gas exiting the nozzle does not heat up, and hence, this type of thruster is called “cold.” However, there are systems that use “warm” gas. These systems have heaters installed, increasing the specific impulse of the thruster [57]. Currently, many models of cold gas thrusters have been operated in space. The flight-operated models of cold gas thrusters are presented in

Table 2. Flight-operated cold gas thrusters.

PS	Entity	Propellant	P, W	T, mN	Isp, s	It, N·s	Size	Mass, kg	Missions (Year)	NORAD ID	Ref.
SNAP-1 ▲	SSTL, Surrey, UK	C4H10	15	46	43	-	<1U	0.5	SNAP-1 (2000)	26386	[66]
MEPSI ▲	The Aerospace Corp., Chantilly, VA, USA	Xe	-	100	30	-	~1U	0.188	STS-11 (2002), STS-11 (2006)	27556, 29647	[59,67]
T3 µPS ▲	TU Delft, Delft, The Netherlands	N2	10	6	>30	-	0.25U	0.2	Delfi-n3Xt (2013)	39428	[51,67]
CNAPS ▲	UTIAS, North York, ON, Canada	SF6	3	12.5…50	45	100	2U	0.26	CanX-4/CanX-5 (2014)	40056, 40055	[51,61,67]
POPSAT-HIP1 ▲	Microspace Rapid, Singapore	Ar	2	0.1…0.3	32	0.6	1U	-	POPSAT-HIP1 (2014)	40028	[62]
NanoProp ● CGP3	GomSpace, Aalborg East, Denmark	C4H10	2	1	60…110	40	0.5U	0.35 **	TW-1 (2015), Astrocast-0301 (2021), Astrocast-0205 (2022), ESTCube-2 (2023)	40928, 54370, 48960	[52,68]
Bevo-2 Cold Gas PS X	Georgia Tech SSDL, Atlanta, GA, USA	R236fa	-	110…150	65…89	58…80	-	0.31 * 0.4 **	Bevo-2 (2015)	41314	[63]
NASA C-POD MiPS ●	VACCO, South El Monte, CA, USA	R236fa	5	10	40	174	0.8U	1.3 **	NanoACE (2017), CPOD A/B (2022)	42844	[69]
MEMS cold gas microthruster ●	CRAS, Hants, UK	N2	<1	1	-	-	0.5U	0.118 *	Ursa Maior (2017)	42776	[70]
MarCO MiPS ●	VACCO, South El Monte, CA, USA	R236fa	0.5	25	42	755	2.5U	3.49 **	MarCO-A/MarCO-B (2018)	43596/43597	[69]
NanoProp ● 6U	GomSpace, Aalborg East, Denmark	C4H10	2	1	60…110	80	200 × 100 × 50 mm3	0.9 **	GOMX-4B (2018)	43196	[71]
I2T5 ●	ThrustMe, Verrieres-le-Buisson, France	I2	5	<0.35	-	75	0.5U	0.9 **	Xiaoxiang 1-08 (2019), NAPA-2 (2021), Robusta-3A (2024)	44706, 48963, 60243	[52]
Cold gas thruster ●	UT Austin, Austin, TX, USA	R236-fa	-	110…170	65…100	-	<1U	-	Armadillo (2019)	44352	[72]
seeker Robotic free flyer propulsion system Ø	NASA, Washington, DC, USA	GN2	-	100	-	-	1.25U	-	Seeker (2019)	44533	[73]
NEA scout propulsion system●	VACCO, South El Monte, CA, USA	R236fa	<55	25	40	500	2U	1.26 * 2.5 **	LiciaCube (2021)	-	[51,70]
ASCENT cold gas PS Ø	Georgia Tech SSDL, Atlanta, GA, USA	-	-	-	-	-	-	-	ASCENT (2021)	51287	[63]
Tianyuan cold gas thruster Ø	NUST, Nanjing, China	-	-	-	-	-	-	-	Tianyuan-1 (2021)	49315	[74]
GDU●	EDB Fakel, Kaliningrad, Russia	N2	9	51.9…96.5	70…120		163 × 95 × 75 mm3	1.1	Geoskan Edelweis (2022)	53385	[75]
ArgoMoon MiPS X	VACCO, South El Monte, CA, USA	R134a	Разогрев 20 Работа 4,3	25	-	72	1.3U	1.43 * 2.07 **	ArgoMoon (2022)	55907	[76]
BioSentinel propulsion system●	Georgia Tech SSDL, Atlanta, GA, USA	R236fa	4	40…70	41…47	79.8	2U	1.28 **	BioSentinel (2022)	55906	[52,77]
OMOTENASHI propulsion system X	VACCO, South El Monte, CA, USA	R236fa	-	25	-	584	1.7U	1.62 *	OMOTENASHI (2022)	99045	[78]
Hamlet▲	NASA ARC, Moffett Field, CA, USA	R236fa	-	2…12	42	-	2U	1.47 * 2.45 **	Starling 6U (2023)	57388	[79]
Politekh Univers-3 Ø	SPbPU, Saint-Petersburg, Russia	R11	-	-	-	-	-	-	Politekh Univers-3 (2023)	57191	[80]
Cold gas propulsion system Ø	GomSpace, Aalborg East, Denmark	-	-	-	-	-	2U	-	Juventas (2024)	-	[81]

* Dry mass. ** Wet mass. ^▲ PS distinct operating parameters, which are available in open information sources. ^● PS demonstrated in orbital flights, but data on its characteristics varies in different sources. ^Ø PS demonstrated in orbital flights, but there is no open data on its characteristics. ^X Propulsion systems that have been launched into space but have failed to successfully demonstrate their functionality in orbital conditions for various reasons.

[57,58,59,60,61,62,63,64,65,66,67,68,69,70,71,72,73,74,75,76,77,78,79,80,81].

Cold gas propulsion systems dominate in formation flying, rendezvous, and low-risk missions primarily due to their inherent advantages in safety, simplicity, reliability, and compatibility with mission requirements. Cold gas systems utilize gases stored at room temperature, eliminating the hazards associated with combustible liquids or corrosive chemicals found in other propulsion technologies. This reduction in hazard mitigates the risk of accidents during handling, integration, launch, and operation phases. Since they do not require ignition or combustion, the likelihood of failures is minimized. Low-risk missions typically prioritize reliability and predictability over raw performance metrics. By minimizing variables introduced by sophisticated hardware, cold gas systems ensure smooth execution of critical maneuvers within predefined margins of error. This makes them ideal candidates for applications where absolute certainty is more important than improvements in efficiency.

Some important notes should be made on propellant selection for the cold gas. In the cold gas, SF₆, C₄H₁₀, R236fa, N₂, and I₂ are mostly used. Its predominant use is driven by several factors such as high storage density, performance characteristics, and relatively known procedures of handling on-ground and during operations [66,67,68,69,70,71,72,73,74,75,76,77,78,79,80,81].

The results presented in Table 2 for the period from 2000 to 2024 reveal similarities in the challenges faced by cold gas thrusters installed on small satellites compared to chemical ones [51,57,67]. Both systems share the characteristic of having a low specific impulse; for example, for cold gas thrusters, I_sp ranges from 30 to 124 s [58,59,61,62,66]. When considering the limited mass of the propellant storage subsystem, this translates into a characteristic speed increment for most missions, amounting to less than 10 m/s over the entire lifetime of the mission [58,62,66]. However, the thrust delivered by these thrusters reaches dozens or even hundreds of mN [59,61,66,69], a level unattainable with electric propulsion systems with similar dimensions. These features render cold gas systems inadequate for tasks requiring significant expenditures of characteristic velocity (major changes in orbit or inclination angle, maintenance of highly elliptic orbits, interplanetary transfers, etc.) but retain their competitiveness in niches such as precise stabilization and orientation correction [57,62,66].

Special attention should be paid to the technological aspects of these systems. Cold gas thrusters can operate either on inert gases (nitrogen, argon, and xenon) [51,59,61,62,67] or active gases (butane, refrigerants, and iodine) [51,52]. Active gases provide higher storage density but show a tendency toward leakage due to their aggressive nature toward seals [63]. This phenomenon has been observed in incidents such as the malfunctions encountered on the OMOTENASHI and Starling SV-1 missions. Additionally, leaks occurred in the MarCO (notwithstanding the successful completion of the mission) and Bevo-2 [63] missions. Despite these leakage risks, active gases continue to be preferred for launches (based on collected statistics over the past 24 years, twice as many thrusters utilizing active gases have been launched) due to their convenient storage characteristics [57,58].

Additionally, it is worth noting a distinct category of cold gas thrusters incorporating MEMS technology [58]. This innovation allows for dry masses of the propellant storage subsystem to be reduced below 0.4 kg, with a typical unit size of 0.5U, while retaining conventional values for specific impulse and total impulse.

Thus, cold gas thrusters, akin to chemical systems, appear to be applicable for executing short-term missions in low Earth orbit with minimal ΔV requirements (orientation correction, de-clustering of satellite formations) [58,62].

3.3. Electric Propulsion

3.3.1. Electrostatic EP

In the electrostatic type of electric propulsion, the energy supplied to the thruster is first used for ionizing the propellant, after which ions are accelerated by an electric field. Electrostatic-type electric propulsion includes Hall-effect (see Figure 6 and Figure 7), ion (see Figure 8 and Figure 9), and electrospray (see Figure 10) thrusters. Such types of propulsion often require a cathode-neutralizer that compensates for the positive charge accumulating on the spacecraft due to the exhaustion of ions. Therefore, developing these thrusters for satellites with a small form factor can become quite complex. Despite this complexity, there is a considerable number of electrostatic-type thrusters that have been flight-operated onboard small satellites. These thrusters are listed in

Table 3. Flight-operated electrostatic-type EP.

PS	Entity	Propellant	P, W	T, mN	Isp, s	It, kN·s	Size	Mass, kg	Missions (Year)	NORAD ID	Ref.
MIPS ● Microwave	UTokyo, Tokyo, Japan	Xe	27	0.21	740	-	340 × 260 × 160 mm3	8.1	HODOYOSHI-4 (2014)	40011	[100]
I-COUPS ● ECR	UTokyo, Tokyo, Japan	Xe	<38	<0.35	1000	-	3U	9.5 **	PROCYON (2014)	40322	[101]
NPT30-I2 ▲	ThrustMe, Verrieres-le-Buisson, France	I2	<65	<2.1	<2500	5.5	96 × 96 × 106 mm3	1.2 **	Hisea-1 (2020), BEIHANGKONGSHI-1 (2020), NorSat-TD (2023)	47297, 46838, 56194	[87,88]
BIT-3 ▲	Busek, Natick, MA, USA	I2	<80	<1.25	<2300	31.7	180 × 88 × 102 mm3	1.5 * 2.9 **	Lunar IceCube (2022), LunaH-Map (2022)	55903	[89,91,93]
Charge Exchange Thruster ▲	University of Sydney, Sydney, Australia	Xe	3	0.027	-	-	100 × 90 × 37 mm3	0.35	i-INSPIRE II (2017), CUAVA-2 (2024)	42731, 60527	[102]
ExoMG-nano ▲	Exotrail, Massy, France	Xe	60	<3	800	<5	2.5U	<2.3 *	M6P (2020), ARTHUR (2021), ELO3 (2023), ELO4 (2023)	44109, 48953, 56216, 56990	[103]
MUSIC-SI ●	Aliena, Singapore	Xe	100	<0.25	<2000	15	1.5U	2 **	NuX-1 (2022)	51073	[104]
MUSIC Hot Mode ●	Aliena, Singapore	Xe	<100	3	1000	15	4U	5 **	ORB-12 Strider (2023)	57483	[104]
NANO ▲	Enpulsion, Schwechat, Austria	In	<40	0.22	3500	<12	0.8U	0.9 ** 0.6 *	Flock 3p (2018), NetSat (2020) NEPTUNO (2021) ***	43119, 46504, 48966	[94,95]
NanoFEEP (GO-2) ●	Morpheus Space, El Segundo, CA, USA	In	<3	0.04	<6000	3.4	90 × 25 × 43 mm3	0.16 * 0.17 **	UWE-4 (2018) ***	43880	[97,98,99]
MICRO R3 ●	Enpulsion, Schwechat, Austria	In	30…120	<1.3	<4500	>5	140 × 120 × 133 mm3	3.9 ** 2.6 *	GMS-T (2021) ***	47346	[94]
NANO AR3 ▲	Enpulsion, Schwechat, Austria	In	45	<0.35	<6000	>5	1U	1.4 ** 1.2 *	AMS (2022), GS-1 (2023) ***	52745, 56372	[105]
S-iEPS	MIT, Cambridge, MA, USA	Ionic liquid	1.5	0.075	<1150	-	96 × 96 × 21 mm3	0.095 *	AeroCube-8 (2015)	41852	[106]
TILE 2 ●	Espace, Hull, CA, USA	Ionic liquid	8	0.05	1800	-	0.5U	0.48	Irvine 01 (2018), Irvine 02 (2018), BeaverCube (2021)	43693, 43789, 53768	[107]
TILE-3 Ø	Accion, Washungton, DC, USA	Ionic liquid	20	0.45	1650	0.755	1U	2 **	D2/AtlaCom-1 (2021)	48922	[108]
Multi-Mode Thruster Ø	Missouri S&T’s Aerospace Plasma Lab, Rolla, MO, USA	Ionic liquid	-	0.25	800	-	-	-	M3 Sat (2024)	-	[109]

* Dry mass. ** Wet mass. *** Missions provided for reference. ^▲ PS distinct operating parameters, which are available in open information sources. ^● PS demonstrated in orbital flights, but data on its characteristics varies in different sources. ^Ø PS demonstrated in orbital flights, but there is no open data on its characteristics.

[82,83,84,85,86,87,88,89,90,91,92,93,94,95,96,97,98,99,100,101,102,103,104,105,106,107,108,109]. Due to the difficulty of downscaling electrostatic-type EP, especially Hall-effect thrusters, most of them are only applicable to satellites sized 6U or larger. However, thanks to extensive expertise with such propulsion units for full-sized spacecraft, some companies have successfully miniaturized them.

Unlike the previously discussed chemical rocket engines and cold gas thrusters, these electrostatic systems exhibit high specific impulse (I_sp = 740–2300 s) and low thrust (approximately 10 mN). This combination allows for significant increases in characteristic velocity with minimal expenditure of propellant. For example, the BIT-3 engine manufactured by BUSEK, Natick, MA, USA reportedly achieves a velocity increment of approximately 2.39 km/s for satellites weighing less than 14 kg, far exceeding the capabilities of analogous cold gas and chemical engines.

One primary challenge facing these systems is scaling. Due to the underlying physics governing their efficiency, they require a minimum length for the acceleration region and a specific geometric ratio for the discharge chamber, physically limiting miniaturization efforts [86]. Consequently, most systems (MIPS, BIT-3, and MUSIC-SI) occupy more than 1.5U of volume and are predominantly suited for satellites in the 6U format and larger.

Similar to cold gas thrusters, developers face the dilemma of selecting either inert or chemically active propellants. Inert gases, such as xenon, offer high acceleration efficiency due to their high molecular mass but necessitate high-pressure tanks and elaborate storage systems [100,103]. Chemical actives, such as iodine (used in systems like NPT30-I2 and BIT-3), provide 30–40% higher storage density but require corrosion-resistant materials for the discharge chamber, seals, valves, etc. [87,88,91]. An alternative approach to chemically active substances emerged with the development of liquid indium-based systems (Enpulsion NANO/MICRO series), where direct evaporation occurs from the surface of the emitter, obviating the need for high-pressure tanks [94,95]. However, these thrusters require preheating to 156 °C to melt the indium, increasing energy consumption during startup [96].

Critical examination of unsuccessful and partially successful missions identifies systemic issues in technology adaptation. The orientation system failure onboard i-INSPIRE-II (2017) resulted from insufficient thrust (27 µN) relative to the satellite’s mass (2.1 kg)—the thrust-to-mass ratio proved inadequate to overcome disturbances in low Earth orbit. Similarly, the limited maneuverability of the system on CUAVA-2 (2024) stems from the absence of a cathode-neutralizer in the charge-exchange architecture, causing positive charging of the spacecraft body and impairing ion acceleration efficiency. These examples underline that established principles effective for larger satellites require fundamental modifications when adapted to the rigid mass-and-volume constraints imposed by CubeSats.

Electrostatic propulsion systems thus hold significant promise for long-duration missions in low Earth orbit (LEO), interplanetary journeys, and orbit correction tasks. However, their low thrust creates challenges in compensating for atmospheric drag in LEO environments, as illustrated by the case of i-INSPIRE-II (2017), where the thrust-to-mass ratio (27 µN for a 2.1 kg satellite) fell short [96]. In parallel, for chemical and cold gas engines, there is an observable trend toward exploring hybrid solutions. One illustrative concept is the Multi-Mode Thruster onboard M³ Sat, which integrates electrostatic acceleration with electrothermal heating [93,109].

3.3.2. Electromagnetic EP

In electromagnetic-type electric propulsion, the ionized propellant is accelerated through effects of electron diamagnetism, extraction of ions by accelerated electrons, or plasma acceleration in crossed electromagnetic fields [110,111]. Such propulsion systems do not require cathode-neutralizers or auxiliary electrodes to function. Electromagnetic-type EP includes electrodeless plasma thrusters (see Figure 11), pulsed plasma thrusters (see Figure 12 and Figure 13), and vacuum arc thrusters. Characteristics of electromagnetic-type electric propulsion for small satellites are presented in

Table 4. Flight-operated electromagnetic-type EP.

PS	Entity	Propellant	P, W	T, mN (Ibit, µN·s)	Isp, s	It, kN·s	Size	Mass, kg	Missions (Year)	NORAD ID	Ref.
Maxwell ●	Phase Four, Hawthorne, CA, USA	Xe	330	5.2	750	-	220 × 120 × 240 mm3	8.4 **	Transporter-1 (2021), Transporter-2 (2021)	48913, 48912	[105,115]
REGULUS-50-I2 ●	T4i, Padova, Italy	I2	50	0.6	<700	3	1.5U	2.5	UniSat (2021), NorthStar Earth&Space (2024)	47945	[113,114,117]
BDEPT ▲	APS, Moscow, Russia	Kr	<120	<10	<1400	1	2U	3.2	HORS 1 (2023), HORS 3 (2024)	57188, 61753	[12]
MPACS Ø	Busek, Natick, MA, USA	PTFE	<5	(80)	827	-	1U	-	FalconSat3 (2007)	30776	[105]
PROITERES ●	Osaka Sangyo University, Osaka, Japan	PTFE	5	(2.47)	340	5 N ● s	100 × 100 × 50 mm3	0.71	PROITERES-1 (2012)	38756	[105]
PPT X	SSTL, Surrey, USA	-	1.5	(0.9)	1340	-	0.25U	-	STRaND-1 (2013)	39090	[118]
PPT ●	Kyushu Institute of Technology, Fukuoka, Japan	PTFE	2.3	(25)	676	-	0.7U	-	Aoba-Velox-III (2016), Aoba-Velox-IV (2019)	41935, 43940	[119]
PPT ●	University of Vienna, Vienna, Austria	PTFE	-	(2.2)	600	5.7	-	-	PEGASUS (2017)	42784	[120]
PPT ●	University of Washington, Seattle, WA, USA	S8	-	-	1200	-	0.6U	-	HuskySat-1 (2019)	45119	[121]
Poseidon M1.5 ●	Miles Space, Tampa, FL, USA	-	1.5	37.5	4800	-	1U	-	Miles (2022)	-	[122]
VERA ●	STAR, Moscow, Russia	POM	5	(30)	620	> 150	83 × 83 × 55 mm3	< 0.5	CUBESX-HSE-2 (2022)	53383	[123]
PETRUS ●	University of Stuttgart, Stuttgart, Germany	PTFE	<1	(10)	699	3.3	84 × 84 × 15 mm3	0.42	GreenCube (2022), SONATE-2 (2024)	53106, 59112	[124]
FPPT ●	CU Aerospace, Boulder, CO, USA	PTFE	48	(240)	3500	5500	1.7U	1.975 * 2.8 **	DUPLEX (2023)	-	[125,126]
µVAT X	University of Illinois in Urbana-Champaign, Urbana, IL, USA	Al	4	(54)	-	-	0.4U	0.15	Illinois Observing NanoSatellite (2006)	-	[127]
μCAT ● GWU	George Washington University, Washington, DC, USA	Ni	<10	(50)	3000	-	0.5U	-	BRICSat-P 2015 CANYVAL-X Tom (2018), BRICSat 2 (2019)	40655, 43136, 44355	[128,129]
XANTUS ●	Benchmark Space Systems, Burlington, VT, USA	Mo	<100	(10)	1764	5000	94 × 94 × 60 mm3	0.85 * 1.4 **	RROCI (2023), RROCI-2 (2024)	55081, 59106	[130]
Neumann Drive ND-15 ●	Neumann Space, Kent Town, Australia	-	<24	(45)	<2000	> 880	150 × 100 × 97 mm3	1.9 **	SpIRIT (2023)	58468	[131]

* Dry mass. ** Wet mass. ^▲ PS distinct operating parameters, which are available in open information sources. ^● PS demonstrated in orbital flights, but data on its characteristics varies in different sources. ^Ø PS demonstrated in orbital flights, but there is no open data on its characteristics. ^X Propulsion systems that have been launched into space but have failed to successfully demonstrate their functionality in orbital conditions for various reasons.

[110,111,112,113,114,115,116,117,118,119,120,121,122,123,124,125,126,127,128,129,130,131].

Unlike electrostatic systems, electromagnetic EP do not require a cathode-neutralizer, simplifying their structure, but they exhibit a wider variation in specific impulse, ranging from 300 s (classic pulsed plasma thrusters) to 4800 s (Poseidon M1.5 thruster).

A distinctive feature of this group of thrusters compared to other types of EP and chemical rocket engines is the possibility of using solid propellants (polytetrafluoroethylene, sulfur, and molybdenum), eliminating the need for complex gas storage and delivery systems [119,120,121,124,130]. This ensures compactness of the propulsion setup; for example, µCAT and PETRUS occupy less than 0.6U volume while providing a specific impulse of 600–3000 s [124,128]. However, the use of solid-fuel systems may be associated with inconsistency in thrust characteristics from activation to activation and limited resource life [119,120]. Successful examples of such thrusters, such as BDEPT (APS), Maxwell (Phase Four), and REGULUS-50-I2 (T4i), confirm the potential of electromagnetic EP for long-duration missions [105,112,113,114].

Electrodeless plasma thrusters deserve particular attention, as they are a subset of electromagnetic EP where the acceleration of the propellant is achieved exclusively through electromagnetic fields, without erosion-prone electrodes in the discharge zone. This architecture eliminates the primary cause of degradation in EP and enables compatibility with a wide range of gaseous propellants, including aggressive components. An example of such a thruster is the BDEPT and MTVEPT thrusters, which were successfully operated on the HORS 1 (2023) and HORS 3 (2024) satellites, respectively. Its unique feature is the capability for multidirectional thrust without movable components. Through rearranging the configuration of electromagnetic fields in the discharge chamber, directional control of the thrust vector in three-dimensional space is achieved. This method reduces the mass of the attitude control system compared to traditional schemes based on flywheels and cold gas thrusters. This improvement allows for a higher attainable increase in characteristic velocity (ΔV) within the limited power output of onboard solar arrays.

Thus, electromagnetic EP occupies a niche in tasks involving precise correction, atmospheric drag compensation, and long-duration missions with moderate requirements for characteristic velocity (ΔV < 500 m/s) on platforms up to 3U format and above. Like electrostatic EP, their application is limited to tasks requiring high thrust (>10 mN) or rapid maneuvers. However, their advantages, such as simplified propellant delivery systems and elimination of the cathode-neutralizer, make them competitive for missions under the tight mass and dimensional constraints of small satellites.

3.3.3. Electrothermal EP

Electrothermal electric propulsion generates thrust by exhausting heated propellant, with heating performed by resistive heaters (resistojets) (see Figure 14), arc discharges (arcjets) (see Figure 15), electromagnetic waves, or laser beams (laser propulsion) (see Figure 16). Typically, inert gases are used as propellants, limiting performance mainly by the maximum wall temperature of the device [100]. Resistojets are best suited for small-satellite applications, given that most arcjets require power exceeding 300 W for proper operation [132,133,134].

Table 5. Flight-operated electrothermal EP.

PS	Entity	Propellant	P, W	T, mN	Isp, s	It, N·s	Size	Mass, kg	Missions (Year)	NORAD ID	Ref.
FMMR X	AFRL, Wright-Patterson Air Force Base, OH, USA	-	2	0.13	80		1U	-	3CS (2004)	43728, 41732	[137]
WARP-DRiVE X	SSTL, Surrey, UK	C4H10	7	-	-	-	0.25U	-	STRaND-1 (2013)	39090	[105]
Arcjet PUC ●	CU Aerospace, Boulder, CO, USA	SO2	15	4.5	68	184	0.25U	0.72 **	8 PS for U.S.A. Air Force (2014)	-	[93,135]
Comet-1000 ●	Bradford Space, Heerle, The Netherlands	H2O	55	17	175	1150	2.3U	1.5 **	HawkEye 360 (2018)	47505	[93,136,138]
AQUARIUS 1U ●	University of Tokyo, Tokyo, Japan	H2O	18	<4	70	<250	1U	1.2 ** 0.8 *	AQT-D (2019), OPTIMAL-1 (2022)	44791, 99207	[139]
AQUARIUS ▲	University of Tokyo, Tokyo, Japan	H2O	<20	<10	<91	-	2.5U	1.3 * 2.5 **	EQUULEUS (2022)	55183	[140]
ARM-A ●	Aurora, Manassas, VA, USA	H2O	<20	<4	100	70	0.3U	0.28 **	AuroraSat-1 (2022)	52427	[90,92]
PBR-10 ▲	Pale Blue, Chiba, Japan	H2O	15	<1	-	<55	0.5U	0.575 **	ArkEdge Space 6U CubeSat	-	[141]
PBR-20 ▲	Pale Blue, Chiba, Japan	H2O	<30	<7	>60	<170	1.25U	1.5 **	SPHERE-1 EYE (2023)	55072	[142]
Steam Thruster 1 ▲	SteamJet Space Systems, Birmingham, UK	H2O	<20	5	172	<100	2U	1 * 1.7 **	PHI-Demo (2023)	57181	[143]

* Dry mass. ** Wet mass. ^▲ PS distinct operating parameters, which are available in open information sources. ^● PS demonstrated in orbital flights, but data on its characteristics varies in different sources. ^X Propulsion systems that have been launched into space but have failed to successfully demonstrate their functionality in orbital conditions for various reasons.

presents the flight-operated electrothermal electric propulsion units [132,133,134,135,136,137,138,139,140,141,142,143].

Electrothermal systems are usually compact (<2U), except for Bradford’s Comet thrusters, which may reach up to 24U in size. They deliver moderate thrust (in the tens of mN) at comparatively low specific impulse (<200 s). Consequently, electrothermal thrusters are commonly employed for short-duration maneuvers, such as final orbit insertion or deorbiting end-of-life satellites.

The electrothermal type of thrusters occupies an intermediate position between EP and cold gas thrusters and chemical rocket engines. They provide thrust in the order of tens of mN, possess compact dimensions (<2U), and simplify the architecture of the propellant storage and delivery system [132,133,134,136]. The specific impulse of such systems does not exceed 175 s, meaning that with a typical propellant mass of 0.5–2.5 kg, the achievable increase in characteristic velocity is limited to values of 5–20 m/s, which is insufficient for long-duration missions with high ΔV requirements [132].

A key trend evident in the development of electothermal EP is the widespread introduction of water as a “green” propellant. Systems such as AQUARIUS (University of Tokyo), ARM-A (Aurora), and PBR (Pale Blue) have shown operational reliability while completely excluding toxic components, simplifying ground testing, and lowering requirements for storage and delivery materials [139,140,141,142]. Particularly noteworthy is the Steam Thruster 1, which, by late 2023, had accumulated over 300 N·s of cumulative impulse across three units onboard PHI-Demo, demonstrating record-level failure-free operation for its class [143].

An intriguing direction in the development of these thrusters, similar to other types, except for electromagnetic ones, is the hybridization of architectures. For instance, the Arcjet PUC thruster (CU Aerospace/VACCO), utilizing microchannel discharges and SO₂ as the working fluid, combines features of resistor jets and arcjets, achieving a specific impulse of 68 s with only 15 watts of power [93,135]. Such solutions help overcome the energy limitations of classical resistor jets without resorting to complex systems incorporating arc discharges, which require powers exceeding 300 W [132,133,134,135,136,137,138,139,140,141,142,143,144].

Thus, electothermal EP occupies a niche in tasks involving short maneuvers with moderate ΔV requirements (<50 m/s), aerodynamic drag compensation in low orbits, end-of-life deorbiting, and spacecraft orientation [132,136,139]. The absence of toxic components marks them as preferable for commercial small-satellite missions amidst tightening environmental regulations for ground operations and space activities.

3.4. Alternative Propulsion

Propulsion systems based on solar sails utilize the effect of light pressure. Photons collide with the reflective surface of the sail, transferring kinetic energy to it. As a result of numerous collisions, the sail moves along the trajectory of the photon flow. The principal scheme of the solar sail propulsion is presented in Figure 17.

The theoretical concept of using a solar sail for movement in space was proposed by Tsander in 1924, although practical implementation of this technology only became feasible at the start of the 21st century. Solar sails do not consume any propellant. However, their thrust characteristics are extremely low, limiting their use to long-distance journeys of lightweight vehicles. For efficient operation, even small vehicles require a substantial sail area to capture sufficient sunlight for propulsion. Furthermore, the orientation of the sail relative to the Sun plays a crucial role. Maximum thrust occurs when the sail faces away from the Sun, aligning itself with the incoming photons. The force exerted on the sail acts perpendicular to its reflective surface, enabling directional control by adjusting the sail’s spatial orientation.

For more efficient conversion of solar pressure into motion, a high reflectivity coefficient of the sail’s working surface is necessary. Additionally, it must exhibit high strength and resilience against external space conditions. Considering the extensive area of the working surface, a framework that combines robustness with light weight is essential. The overall design of the propulsion system must inherently include folding capabilities because deploying an unfurled solar sail directly into orbit would be impractical given its substantial dimensions. Designing the thrust mechanism and engineering suitable materials for the active surface present significant challenges. Consequently, it was not until recent times that constructing and testing solar sails became viable within Earth’s orbit. In

Table 6. Flight-operated solar sails onboard small satellites.

PS	Size	Deployment	Area, m2	Mass, kg	Material	Launch Date	Destination	NORAD ID	Ref.
IKAROS	Square, 14 m × 14 m	Spinning	196	310	PA	21 May 2010	Venus	36577	[144,145]
NanoSail-D2	Square, 3.75 m × 3.75 m	Spinning	14	4	PET/Al	20 November 2010	LEO	37361	[146,147]
LightSail 2	Square, 5.6 m × 5.6 m	Spinning	32.6	4.93	PET/Al	25 June 2019	LEO	44420	[148,149]
NEA-Scout	Square	Deploying rod	86	12	PET/Al	16 November 2022	LEO	-	[150,151]
ACS3	Square, 9.9 m × 9.9 m	Deploying rod	80	9	PET/Al	23 April 2024	LEO	59588	[152,153]

, characteristics of the flight-operated solar sails are presented [144,145,146,147,148,149,150,151,152,153].

Alternative propulsion systems considered in this section differ markedly from previously reviewed systems, as they entirely eliminate the need for propellant by harnessing photon pressure from solar radiation to accelerate the spacecraft [144,145,146,147,148,149,150,151,152,153]. However, this approach leads to drastically lower thrust generation, orders of magnitude weaker even than the weakest electrostatic EP. Consequently, solar sails prove effective solely in long-duration missions > 3 years, where accumulated ΔV can reach hundreds of meters per second without consuming mass [148,150,152].

To achieve ΔV comparable to EP, approximately 1 km/s, a sail area exceeding 100 m² is required, with the satellite mass kept below 10 kg. This yields a fragile structure that is highly susceptible to micrometeoroid erosion and necessitates complex deployment mechanisms, as evidenced by partial failures in missions such as NEA Scout (2022) [150,151].

Promising directions for developing such systems lie in pursuing hybrid solutions with EP. For instance, coupling a solar sail with a water-based electothermal resistojet (as proposed in the ACS3 mission concept) permits the sail to perform primary acceleration while the electric motor handles precise trajectory correction maneuvers [152,153]. This approach compensates for the inability to move opposite to the solar flux direction and enhances maneuverability through minimal propellant reserves.

4. Discussion

The discussion begins by reviewing the operational characteristics of propulsion systems installed aboard small-form-factor satellites, particularly those used in CubeSats. Subsequently, various criteria are examined to evaluate current propulsion technologies, identify prevailing trends among operational systems, and forecast potential advancements in next-generation propulsion solutions.

4.1. Use of Different Propulsion System Types Aboard Civil Small Satellites

Figure 18 illustrates the normalized frequency of small-satellite missions employing diverse propulsion systems between 2000 and 2024. Over this period, there has been a notable rise in the number of small satellites utilizing propulsion technologies. This indicates an increasing focus on controllable small-satellite missions. Presently, such missions primarily involve communications and remote sensing applications. Future projections suggest continued growth in highly maneuverable small satellites, enabling services like orbital maintenance and contributing to resilient distributed space architectures.

Figure 19 displays a pie chart showing the distribution of small-satellite missions that employed various propulsion systems. It emphasizes the predominance of electric propulsion technology, especially electrospray propulsion, accounting for 82% out of the 305 missions analyzed. These findings highlight the rising adoption of electric propulsion systems in small-satellite missions, attributed to their enhanced efficiency and extended service-free operational life spans [153].

Figure 20 examines the distribution of chemical propulsion methods. Among civilian applications, cold gas propulsion prevails over other chemical options—monopropellants and bi-propellants follow closely behind. Its popularity stems from its inherent simplicity and dependability, making cold gas thrusters ideal for compact spacecraft designs, where minimizing mass and reducing complexity are crucial considerations.

Figure 21 offers a comprehensive analysis of electric propulsion systems deployed in small-satellite missions, identifying electrospray propulsion as the leading option, complemented by PPT and resistojet systems. This insight underlines the ascendance of electric propulsion technologies, favored for their exceptional energy efficiency and extended functional lifespan, and thus proving indispensable for sustained deep-space exploration efforts [153,154,155,156,157].

Figure 18, Figure 19, Figure 20 and Figure 21 together illustrate the dynamic evolution of small-satellite propulsion systems, showcasing the growing prominence of electric propulsion, notably electrospray propulsion, in contemporary missions. This transition mirrors the escalating need for higher-efficiency and longer-lasting systems capable of sustaining the expanding intricacy and mission durations associated with distributed space networks reliant upon small satellites.

4.2. Evaluation Criteria

To facilitate a more in-depth examination of trends and prospects in small-satellite propulsion systems, an evaluation framework is proposed. This framework employs a tailored scoring system to evaluate multiple facets, including physics understanding, developmental complexity, operational reliability, material accessibility, production simplicity, scientific outreach, and educational opportunities. Numerical ratings ranging from +1 to −1 are allocated accordingly, denoting favorable, neutral, or adverse qualities, respectively. A detailed explanation of these evaluation metrics is outlined in

Table 7. Explanation of scores used.

Score	Description
+1	Highly understood physics processes, simplicity of development and testing, operational confidence, materials availability, ease of production processes, scientifically widespread, and educationally widely available
0	Partially understood physics processes, relatively complex development and testing, operational uncertainty, limited material availability, challenging production processes, limited scientific dissemination, and limited educational accessibility
−1	Poorly understood physics processes, significantly complex development and testing, persisting operational challenges, scarce material availability, challenging production processes, restricted scientific dissemination, and restricted educational accessibility

.

4.3. Physics Rationale

The decision to choose a particular thruster system for a space mission can be based on evaluating the physics rationale based on three critical factors: the accurate understanding of physical processes, the detailed physics-based descriptions, and the rigorous testing validations. Each type of thruster offers distinct advantages and challenges. The summary of the discussion on the physics rationale for each type of propulsion systems based on the following criteria: process understanding, physics description, testing validation is presented in

Table 8. Physics rationale of trends in propulsion technologies for small satellites.

Propulsion Type	Processes Understanding	Physics Description	Testing Validation	Overall Score
Monopropellant rocket	1	1	1	3
Bi-propellant rocket	1	1	1	3
Hybrid rocket	0	0	1	1
Cold gas	1	1	1	3
HET	0	1	1	2
IT	1	1	1	1
Electrospray	1	1	1	3
PPT	0	0	0	0
VAT	0	0	0	0
EPT	0	0	1	1
Resistojet	1	1	1	3
Arcjet	1	1	1	3
Laser propulsion	1	1	1	3
Solar sail	1	1	1	3

.

Physics understanding is not typically problematic for small-satellite chemical propulsion systems. Monopropellant/bi-propellant engines benefit from fully understood thrust mechanisms. Hybrid engines combine solid fuels with liquid oxidizers, offering flexible design possibilities, but face scaling challenges for miniature satellites. Cold gas for thrust generation utilizes pressurized gas that exhausts from the nozzle. Hall-effect thrusters are one of the best studied EP. Nonetheless, their requirements for dimensional ratios between their structural elements represent a constraint for the development of missions based on satellites with a small form factor [158]. The processes of the ion thrusters are also well-studied, but the development of these systems for small satellites requires the use of expensive materials and technological processing [159]. Pulsed plasma thrusters suffer from limited physics understanding of responsiveness [160]. Vacuum arc thrusters are also poorly understood from a physics point of view [161]. The key challenge for EPT is in understanding the processes of thrust generation since there are no well-established theories for their description [22].

The physics of processes in the resistojets is fully understood [162]. Arcjets’ physics of operation is fully understood [163]. The physics of the processes of laser propulsion is relatively well-understood, and there are several successful realizations of laser propulsion that can be operated onboard small satellites [164]. Solar sails exploit photon reflection off reflective surfaces to harness radiation pressure. These processes are well-studied, but the main limitation for these systems is the on-ground testing.

4.4. Engineering Rationale

When determining the appropriateness of different types of thrusters for modern small-satellite missions from an engineering viewpoint, four key criteria must be taken into account: developmental effort, testing effort, integration effort, and the capacity to realize thrust vectoring.

Table 9. Engineering rationale of trends in propulsion technologies for small satellites.

Propulsion Type	Development Effort	Testing Effort	Integration Effort	TVC Realization	Overall Score
Monopropellant rocket	1	0	0	1	2
Bi-propellant rocket	1	0	0	1	2
Hybrid rocket	0	0	0	1	2
Cold gas	1	1	1	1	4
HET	0	1	0	0	1
IT	0	1	0	0	1
Electrospray	0	1	1	1	3
PPT	0	0	−1	0	−1
VAT	0	0	−1	−1	−2
EPT	0	1	1	1	3
Resistojet	1	0	0	1	2
Arcjet	1	0	0	1	2
Laser propulsion	0	1	1	1	3
Solar sail	1	−1	0	1	1

outlines a comparative analysis of each thruster type based on these factors.

Monopropellant rocket engines, complicated propellant formulas, or specialized hardware might lead to added engineering [165]. Bi-propellant rocket engines carry a high testing load because of complex fuel interactions, necessitating sophisticated simulations and expertise in fluid dynamics and thermochemistry [166]. Hybrid rocket engines have medium–high testing complexity, attributed to materials compatibility and ignition mechanisms introducing design challenges. Cold gas thrusters have low development effort, as their simple design reduces the requirements for testing facilities [167]. Miniaturized Hall-effect thrusters face high engineering barriers, driven by specialized magnets and electronics, increasing design challenges [168]. Miniaturized ion thrusters represent the peak of engineering difficulty, as advanced ion-optical systems and cathode-neutralizers demand intensive computational modeling and simulation before laboratory prototypes [169]. Electrospray thrusters have design challenges centered mostly on emitter array geometry and power circuits [170]. Pulsed plasma thrusters bear medium–high development costs, caused by achieving the wanted mode of operation—electromagnetic or thermal modes of operation—and eliminating electrode wear, posing significant engineering obstacles. In addition, PPT poses challenges with its integration into satellites of small form factor [171,172,173]. Vacuum arc thrusters have problems with managing metal vapor dynamics and cathode erosion [174]. Electrodeless plasma thrusters have engineering obstacles with the radiofrequency coupling efficient realization [175]. Resistojets have difficulties with the heating element design [176]. Arcjets face engineering obstacles related to the electrode erosion rates, nozzle throat geometry, and power regulation schemes [177]. Laser propulsion may have high engineering costs, but the overall developmental process is not complicated [178]. Solar sails claim high engineering complexity, requiring optimized reflector geometries and deployment mechanisms [179].

Addressing thrust-vectoring realization, monopropellant, bi-propellant, hybrid rocket engines, cold gas, resisitojets, and arcjets demonstrate good capacity for thrust-vectoring realization, as they have dense exhausting flows. Hall-effect and ion thrusters are practically costly, from the commercial and engineering points of view, and have opportunities for the realization of thrust-vectoring. Electrospray thrusters have a great opportunity for thrust-vectoring; for example, by implementing a sectioned emitter or using multi-emitter arrays, affording independent control over the thrust direction. Pulsed plasma thrusters have been demonstrated in orbit to have the capability for controlling the thrust vector direction by implementing the geometric approach for its realization. Electrodeless plasma thrusters have also been demonstrated to control the thrust vector direction in-orbit. Laser propulsion can realize the capability to control the thrust vector direction in several ways [12]. Solar sails’ structure can enable dynamic sailing feats, but only in the proximity to stars.

4.5. Operational Rationale

When evaluating propulsion systems for small-satellite missions from an operational standpoint, three critical attributes must be assessed: specific mass/volume per unit of thrust and specific impulse, specific impulse, and thrust adjustability. In

Table 10. Operational rationale of trends in propulsion technologies for small satellites.

Propulsion Type	Specific Mass/Volume	Specific Impulse	Thrust Adjustability	Overall Score
Monopropellant rocket	1	−1	0	0
Bi-propellant rocket	1	−1	0	0
Hybrid rocket	1	−1	1	1
Cold gas	1	−1	1	1
HET	0	0	1	1
IT	1	1	0	2
Electrospray	0	1	1	2
PPT	0	0	0	0
VAT	0	0	0	0
EPT	1	0	1	2
Resistojet	1	0	1	2
Arcjet	1	0	1	2
Laser propulsion	1	1	0	2
Solar sail	−1	0	−1	−2

, the comparison between different propulsion systems based on the assessment criteria proposed is presented.

It should be noted that for small-form-factor satellites, featuring limited volume and power availability, the specific impulse is one of the key characteristics that affect the maneuvers that the propulsion system can assist a spacecraft with. This characteristic directly affects all propulsion capabilities. The subsequent analysis relies upon information presented in Section 3 and mainly focuses on the discussion of the specific impulse of different propulsion technologies. Monopropellant rocket engines demonstrate satisfactory performance, characterized by a specific impulse of approximately 280 s, enabling fundamental maneuvering capacities. These engines display moderate modular properties and possess the ability to adjust power output within predetermined boundaries. Bi-propellant rocket engines yield superior specific impulse, varying between 300 and 450 s, facilitating broader operational ranges. This type of engine exhibits expansive throttling capabilities, thereby increasing its adaptability for dynamic space missions. Hybrid rocket engines have an I_sp of 280–300 s, which is comparable to other chemical propulsion systems. However, they may encounter limitations in flexibility due to rigid output configurations. Cold gas thrusters underperform significantly among other propulsion systems, exhibiting a specific impulse of merely 100 s, which restricts their application predominantly to precise pointing tasks. Hall-effect thrusters generate flows with specific impulses that range from 500 to 1500 s for compact designs. Consequently, these systems prove advantageous for deep-space explorations, despite having limited adjustment scope. Ion thrusters surpass all other systems, boasting an exceptionally high specific impulse above 3000 s. Electrospray thrusters stand out with specific impulse levels extending from 1000 to 5000 s, providing unequaled accuracy via independently adjustable emitters. Pulsed plasma thrusters deliver mediocre results, achieving specific impulse values ranging from 100 to 300 s. Although modifications are possible through varied frequencies, such changes lead only to tolerable outcomes rather than remarkable achievements. Vacuum arc thrusters function suboptimally, producing a specific impulse of about 1200 s. Electrodeless plasma thrusters correspond closely to Hall-effect thrusters, considering compact design performance, offering acceptable specific impulse rates between 500 and 1200 s. Resistojets and arcjets fall short in comparison, yielding specific impulse outputs confined to the range of 100–300 s. Their functionality resembles binary switches, lacking subtle control mechanisms.

4.6. Production Rationale

From a production perspective, the selection of propulsion systems for small-satellite missions pivots around three core criteria: material availability, production processes, and storage simplicity. The summary of the discussion on the production rationale for each type of propulsion system based on the following proposed criteria for this rationale is presented in

Table 11. Production rationale of trends in propulsion technologies for small satellites.

Propulsion Type	Materials Availability	Processes of Production	Storage Simplicity	Overall Score
Monopropellant rocket	1	1	0	2
Bi-propellant rocket	1	1	0	2
Hybrid rocket	1	0	0	1
Cold gas	1	1	0	2
HET	0	0	0	0
IT	0	0	0	0
Electrospray	1	0	1	2
PPT	1	0	1	2
VAT	1	0	1	2
EPT	1	1	0	2
Resistojet	1	1	0	2
Arcjet	1	1	0	2
Laser propulsion	1	0	0	1
Solar sail	1	0	1	2

.

Material availability plays a crucial role in determining manufacturability. Monopropellant rocket engines utilize commonly accessible chemicals such as hydrazine and hydrogen peroxide. Hybrid rocket engines employ readily obtainable materials like paraffin wax and nitrous oxide. Cold gas, Hall-effect, ion, and electrodeless plasma thrusters have the advantage of being able to use widely available gases, including nitrogen and argon. The complexity involved in production varies considerably across different types of propulsion systems. For instance, monopropellant rocket engines depend on conventional manufacturing techniques, such as machining and welding, whereas bi-propellant rocket engines necessitate sophisticated fabrication processes along with rigorous sealing verifications. Regarding storage requirements, there exist notable differences among various systems. While cold gas thrusters allow safe storage in pressurized containers, bi-propellant rocket engines might entail demanding cryogenic storage conditions. Ultimately, propulsion systems employing solid-state propellants and incorporating minimal electronics excel in terms of material accessibility, simplified manufacturing procedures, and reduced storage demands when compared to more advanced alternatives [180].

4.7. Societal Rationale

From a societal perspective, the selection of thruster systems for small-satellite missions rests on three essential criteria: historical legacy, scientific prevalence, and educational outreach. Historical legacy refers to the chronology and maturity of the technology. Scientific prevalence signifies the pervasiveness of the technology studied and developed across the world. Educational accessibility relates to how widespread the educational programs that cover the study and development of the distinct technology are. The societal rationale assessment is presented in

Table 12. Societal rationale of trends in propulsion technologies for small satellites.

Propulsion Type	Historical Legacy	Scientific Prevalence	Educational Outreach	Overall Score
Monopropellant rocket	1	1	1	3
Bi-propellant rocket	1	1	1	3
Hybrid rocket	0	0	0	0
Cold gas	1	1	1	3
HET	1	1	1	3
IT	1	1	1	3
Electrospray	0	0	0	0
PPT	1	1	1	3
VAT	0	0	0	0
EPT	1	0	0	1
Resistojet	1	1	1	3
Arcjet	1	1	1	3
Laser propulsion	0	0	0	0
Solar sail	0	0	0	0

.

Monopropellant rocket engines and bi-propellant rocket engines trace their lineage back to early space pioneers, establishing reliable reputations for key missions. They enjoy widespread scientific dissemination and educational accessibility. Hybrid rocket engines represent an innovative shift, rising in importance recently, and are accompanied by noteworthy accomplishments. Cold gas thrusters originate from earlier eras, favored for their simplicity and reliability. These propulsion systems share similarities with monopropellant and bi-propellant rockets in terms of scientific prevalence and educational availability. Hall-effect thrusters emerged in the Soviet Union during the 1960s, later spreading globally and becoming integral components in numerous exploration missions. Among electric propulsion technologies, Hall-effect thrusters hold the longest developmental and operational history. Physicist Morozov initiated investigations into applying the Hall effect to propulsion systems in the mid-twentieth century, leading to prototype creation in the 1950s. Development intensified in the 1960s, with successful space testing occurring in the 1970s. Nowadays, Hall-effect thrusters serve as vital tools in space exploration, powering both communication satellites and deep-space probes. Ion thrusters experienced gradual evolution prior to their demonstration aboard NASA’s Deep Space 1 mission. Electrospray thrusters constitute emerging players garnering influence within academic communities. Despite this, the space propulsion sector has provided numerous units to support small-satellite missions. Pulsed plasma thrusters arose in the 1960s, experiencing cautious resurgence in contemporary times, undergoing active terrestrial study and space-based experimentation. Historically, vacuum arc thrusters remained largely dormant before recent activation beyond laboratory confines and transition towards flight operation tests. Electrodeless plasma thrusters stemmed from radiofrequency plasma sciences, gradually accumulating supporters. Thrust-vectoring thruster developments based on this technology have reached continuous space operations. Resistojets and Arcjets derive from older methodologies and continue to undergo steady investigation and refinement. Laser propulsion is being developed continuously. Remarkably, laser propulsion intended for small satellites once approached initial flight testing, aborted by an ill-fated space vehicle malfunction, causing the mission’s demise in the Atlantic Ocean. Solar sails have fascinated humanity for centuries, transforming into practical realities in modern times.

To summarize, monopropellant rockets, bi-propellant rockets, cold gas, Hall-effect, ion, pulsed plasma thrusters, resistojets, and arcjets dominate current practices, sustained by robust historical foundations, ubiquitous scholarship, and inclusive teaching. Particularly, monopropellant and bi-propellant rockets, Hall-effect, ion, and electrospray thrusters attract substantial investment and are readily accessible for study. Alternatively positioned technologies such as vacuum arc thrusters, laser propulsion, and solar sails struggle incrementally, hindered by obscurity. The scientific community, pursuing innovation, focuses increasingly on hybrid propulsion systems and electrosprays, uncovering promising opportunities.

4.8. Trends and Perspectives in Propulsion Systems for Small Satellites

The overall scores based on multiple rationales for small-satellite propulsion technologies are presented in

Table 13. Overall development and rationales for propulsion technologies for small satellites.

Propulsion Type	Overall Score
Monopropellant rocket	10
Bi-propellant rocket	10
Hybrid rocket	5
Cold gas	13
HET	7
IT	7
Electrospray	10
PPT	4
VAT	0
EPT	9
Resistojet	12
Arcjet	12
Laser propulsion	9
Solar sail	4

.

Table 7 indicates that cold gas propulsion and electrosprays lead in overall development and utilization rationale among chemical and electric propulsion systems used in small satellites. Their popularity arises mainly because they are simple, reliable, and suitable for very small-sized satellites. This conclusion matches the data showing predominance of certain propulsion technologies in chemical (Figure 20) and electric (Figure 21) propulsion.

Other important electric propulsion candidates include resistojets and arcjets, which score highly. They offer accurate control over trajectory corrections and efficient energy consumption. However, according to Section 4.1, these technologies are not frequently adopted on small satellites. One explanation could be that they generate excess heat, requiring extra cooling systems, and making them impractical for small satellites.

Certain propulsion technologies lag far behind others. As shown in Table 7, solar sails earn only four points, highlighting their limited usefulness given current constraints and difficulties in deploying them. Pulsed plasma thrusters similarly score low, indicating problems with integration and effectiveness on small satellites. Vacuum arc thrusters perform the worst, receiving zero points, revealing major technical issues.

Emerging technologies show future potential despite initially low scores. Examples include hybrid rockets, planar Hall-effect thrusters, and laser-based propulsion. Hybrid rockets merge solid fuels with liquid propellants, combining the benefits of each approach. Planar Hall-effect thrusters can fit into compact volumes while maintaining high performance. Laser-driven systems efficiently convert light energy into motion, potentially surpassing conventional propulsion methods in terms of thrust efficiency.

Interestingly, there are no distinctive differences between the propulsion systems for commercial use and for scientific operations. The only two differences are that for the commercial operations, the industry primarily uses the propulsion systems with the highest flight heritage, for example, Hall-effect thrusters, and is looking for the use of a low-cost propellant substance, such as argon, considering EP. Governmental missions are not considered in this review, as they are dual-use missions. Looking ahead, several clear trajectories emerge concerning propulsion development pathways for small satellites. Ongoing refinements will focus heavily on shifting space missions from static to dynamic ones, requiring propulsion systems that can significantly increase the maneuverability of satellites. The development of such propulsion technologies requires the reconsideration of conventional schemes and the combination of several propulsion technologies in a single device. A promising direction for the development of propulsion systems on small satellites is the creation of hybrid systems that integrate multiple types of engines into a single module to cover a wide range of maneuvering tasks within a single mission. This necessitates developments aimed at increasing the energy efficiency of engines to enable their combination under strict energy capacity constraints.